The present invention relates generally to gas turbine engines, and, more specifically, to cooling hot components therein.
Advancements in aircraft gas turbine engines have improved both fuel efficiency and core specific power. These advances have occurred in large part by the improved thermodynamic cycle achieved by simultaneous increases in the overall pressure ratio and in the turbine inlet temperatures. A major step forward was achieved by the introduction of air cooling for the critical hot components such as the combustor, and turbine blades, nozzles, shafts, and disks. The rate of further advances has been paced by both the technology for improving air cooling and the availability of high temperature materials for the critical components. These advances slowed as the physical limits of metallic materials were approached.
The current method of air cooling utilizes a portion of the compressor exit air as the cooling source specifically bled and channeled for cooling the hot components. Such bleed air is parasitic since it is not used in the combustion process for generating power. As the pressure ratio is ever increased in turbine designs, the cooling air temperature correspondingly increases and therefore the quantity of such air must be increased. Similarly, increases in turbine inlet temperature also result in corresponding need to increase the cooling air. These increases of parasitic cooling air eventually limit the cycle gains and no further improvements are possible unless new methods are utilized. Methods which attempt to cool the cooling air discharged from the compressor have been limited by the attendant pressure losses encountered in the cooling process. In some conventional designs, a separate compressor is provided to repressurize the cooling air, with a resultant increase in complexity of the engine and power loss thereby.